The present invention relates to the field of axial flow machine cooling or ventilation systems, and, more particularly, although not exclusively, to the field of gas turbine engine component cooling.
With reference to FIG. 1, a ducted fan gas turbine engine according to the prior art generally indicated at 10 has a principal and rotational axis 11. The engine 10 comprises, in axial flow series, an air intake 12, a propulsive fan 13, an intermediate pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, and intermediate pressure turbine 18, a low-pressure turbine 19 and a core engine exhaust nozzle 20. A nacelle 21 generally surrounds the engine 10 and defines the intake 12, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 12 is accelerated by the fan 13 to produce two air flows: a first air flow into the intermediate pressure compressor 14 and a second air flow which passes through a bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 14 compresses the air flow directed into it before delivering that air to the high pressure compressor 15 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive, the high, intermediate and low-pressure turbines 17, 18, 19 before being exhausted through the nozzle 20 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines 17, 18, 19 respectively drive the high and intermediate pressure compressors 15, 14 and the fan 13 by suitable interconnecting shafts.
Alternative gas turbine engine arrangements may comprise a two, as opposed to three, shaft arrangement and/or may provide for different bypass ratios. Other configurations known to the skilled person include open rotor designs, such as turboprop engines, and turbojets, in which the bypass duct is removed such that all air flow passes through the core engine. The various available gas turbine engine configurations are typically adapted to suit an intended operation which may include aerospace, marine, power generation amongst other propulsion or industrial pumping applications.
When a gas turbine engine as shown in FIG. 1 is in operation many of the components of the engine, in particular those components in the high pressure sections immediately downstream of the combustor 16, experience temperatures which are in excess of 1500° C. Such high operating temperatures are often many hundreds of degrees greater than the actual melting points of the individual components, and it is therefore necessary to provide such components with a supply of coolant, such as air. It is generally desirable for the high pressure turbine to withstand as high a combustion exhaust temperature as possible, as this results in an increase in the level of thrust that the engine provides. Thus, there is a need to cool the high pressure turbine and other components immediately downstream of the compression system and combustor as efficiently as possible.
A typical way of providing a coolant duct for the high pressure section of a gas turbine engine is to provide a ventilation cavity located between the high pressure compressor and the high pressure turbine, which may then act as part of a cooling system for the high pressure turbine. Air is extracted from the boundary layer of the main gas path from the high pressure compressor, and fed to the high pressure turbine along the cavity. The air passing through the high pressure compressor has passed through multiple compressor stages and is thus relatively hot compared to ambient temperature, despite being cooler than the combustion exhaust temperature. The air which is extracted from the boundary layer has a higher temperature than the main gas path of the high pressure compressor, and this can have a detrimental impact on engine components both in the cavity, and in the high pressure turbine.
It is an aim of the present invention to provide a system in which an improved cooling flow regime is established so as to reduce thermal loading on the high pressure section of the engine.